A turbine exhaust case (TEC) includes an outer case, an inner case, and a plurality of struts structurally connecting the inner case to the outer case. At least one of the struts has an airfoil body with a hollow core. A leading edge stiffener is provided at the radially inner end of the airfoil body. The leading edge stiffener projects into the hollow core and merges with a stiffener ring projecting from a radially inner surface of the inner case.

An internal core profile for a second stage turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

Turbine blades with non-axisymmetric forward features are provided. A turbine blade may include a platform and an airfoil extending radially outward from the platform and configured to extend into a fluid flowpath. The airfoil separates an upstream portion of the fluid flowpath from a downstream portion of the fluid flowpath. A sealing member extends axially from the platform toward a stationary nozzle adjacent the platform and separates the fluid flowpath from a wheel space. The platform may have a forward face between the sealing member and the airfoil and, optionally, a forward axial face between the forward face and the airfoil. The forward face or forward axial face may face the upstream portion of the fluid flowpath and may have a profile that is non-axisymmetric with respect to its centerline axis.

A stator vane segment for a turbomachine is provided, in particular for a gas turbine, in particular for a turbine stage of a gas turbine. The stator vane segment has at least one stator vane and at least one shroud, in particular an outer shroud, having at least one first profile disposed on the shroud and adapted for attachment of the stator vane segment to the turbomachine casing, the profile extending in the circumferential direction at least partially over a circumferential length of the stator vane segment along the shroud of the stator vane segment and having at least one functional surface which extends at least partially in the axial direction and in the circumferential direction, at least one functional surface of at least one profile having at least two different curvatures in the circumferential direction in at least one radial plane perpendicular to an axis of rotation of the turbomachine in at least one temperature range below a defined operating temperature of the turbomachine.

A waste heat recovery system for recovering waste heat of in internal combustion engine includes a turbine expander. The turbine expander includes a turbine blade, a shaft coupled to and rotatable by the turbine blade, and a nozzle assembly. The nozzle assembly includes a nozzle block disposed about the shaft and adjacent the turbine blade, a first nozzle component coupled to the nozzle block, and a second nozzle component coupled to the nozzle block. The first nozzle component defines a first nozzle having a first geometrical configuration. The second nozzle component defines a second nozzle having a second geometrical configuration that is different from the first geometrical configuration. The waste heat recovery system also includes a flow control device in fluid communication with the turbine expander. The waste heat recovery system further includes a controller in communication with the flow control device.

A turbine nozzle for a gas turbine engine includes a plurality of nozzle segments that are configured to be assembled into a full ring such that each one of the plurality of nozzle segments is adjacent to another one of the plurality of nozzle segments. Each one of the plurality of nozzle segments includes an endwall segment and a nozzle vane. The turbine nozzle includes a feather seal interface defined by endwall segments of adjacent ones of the plurality of nozzle segments. The feather seal interface is defined along an area of reduced pressure drop through a pressure field defined between adjacent nozzle vanes of the plurality of nozzle segments to reduce leakage through the plurality of nozzle segments. The turbine nozzle includes a feather seal received within the feather seal interface that cooperates with the feather seal interface to reduce leakage through the plurality of nozzle segments.

A vane has an airfoil extending between a radially outer platform and a radially inner platform. At least one of the platforms has nominally radially thinner portions, and a pad defining a radially thicker portion. The pad has a radial thickness that is greater than a thickness of the nominal radially thinner portions. The pad surrounds an outer periphery of the airfoil on a side of the radially outer platform. The pad has a varying radial thickness. A mid-turbine frame and a gas turbine engine are also disclosed.

A nozzle assembly is provided which is, in part, formed of a low coefficient of thermal expansion material. The assembly includes a nozzle fairing formed of the low coefficient of thermal expansion material and includes a metallic strut extending radially through the nozzle fairing. Load is transferred from the nozzle fairing to a static structure in either of two ways: first, the strut may receive the load directly and/or second, load may be transferred from the nozzle fairing to at least one of the inner and outer support rings. Further, the nozzle fairing and strut may allow for internal airflow for cooling.

A turbocharger has a turbine housing scroll that is meridionally divided into first and second scrolls. A nozzle ring disposed in the turbine nozzle has an array of circumferentially spaced first vanes and an array of circumferentially spaced second vanes. The first vane passages between first vanes are configured to blow exhaust gas onto the turbine blade leading edges. Likewise, the second vane passages are configured to blow exhaust gas on the turbine blade leading edges. The first and second vanes are circumferentially staggered relative to each other such that the turbine blade leading edges receive exhaust gas from the first and second blade passages in interleaved fashion about a circumference of the turbine wheel.

A nozzle for a turbine system includes an airfoil, an inner sidewall, and an outer sidewall. Each of the inner sidewall and outer sidewall each includes a peripheral edge defining a pressure side slash face, a suction side slash face, a leading edge face, and a trailing edge face. At least one of the inner sidewall pressure side slash face, the inner sidewall suction side slash face, the outer sidewall pressure side slash face, or the outer sidewall suction side slash face includes a first swept surface extending at a first angle relative to a nominal slash face angle and a second swept surface extending at a second angle relative to the nominal slash face angle. The first and second swept surfaces meet at a joining line that is circumferentially aligned with a stiffening member extending circumferentially on a respective sidewall.